Method to control thickness in composite parts cured on closed angle tool

ABSTRACT

The present invention provides a method for controlling the thickness of composite laminates cured on closed angel tools. The invention utilizes a peel ply rather than a breather during the preparation for cure of laminate parts having a substantially uniform initial thickness. This method allows composite parts to be formed with greater uniformity and without the need for extra hardware. The composite part formed by the above method may find use in a wide variety of applications, including, for example, automotive and aerospace applications. Thus, for example, a composite part formed in accordance with the present invention is ideally suited for use as a shear tie in a commercial aircraft, which are used to secure the inner framework of the aircraft to the airplane skin.

TECHNICAL FIELD

The present invention generally relates to preparation of shapedcomposite materials and more specifically to a method for controllingthe thickness of composite parts formed on a closed angle tool.

BACKGROUND ART

Cured laminate composite parts may conventionally be formed frompreimpregnated (“prepreg”) tape of epoxy and carbon or glass fiber. Theprepreg plies are shaped (i.e. bent or curved) and applied to a femalecuring tool. Conventional processing on female cure tools produces outof tolerance laminate thickness conditions with radii that are thickerthan nominal regions. Excessive thickening in the radius beyond theallowable thickness tolerance produces two main problems. First, therewill either be large resin pockets between a few plies or the resinlayers between many plies will be thicker than normal. Second,fasteners, which need to be located near the inner radius of the sheartie, will be hindered if the thickness of the radius region encroacheson the web or the flange. Either scenario may result in the compromiseof mechanical properties or assembly fit-up.

Thus, recently, there have been methods developed to control withintolerance the thickness of the entire cured laminate, including the areathrough the radii. For instance, one method, known as radius pressureintensifying (“RPI”), has been proposed. RPI is a process in which acorner block is coupled against the inner radius of an uncured laminatematerial. A pressurized bladder is introduced between the block and alarger heated tooling and inflated to a desired pressure essentiallypreconsolidating the radius under heat and pressure. The RPI is thenremoved from the part and the laminate part is then bag finished andcured as usual. While RPI can achieve parts having substantially uniformthicknesses, there are disadvantages in utilizing the RPI process. Forexample, RPI requires unique tooling, the need for a pressurized airsupply, and a heated cure tool in order to work. Further, if thelaminate part is cycled multiple times utilizing the RPI method, thethickness of the inner radii may be thinned, thus affecting theuniformity of thickness of the part.

Another method for uniformly controlling the thickness utilizes variousforms of soft or hard cauls to control the thickness through the radii.However, the use of cauls may cause bulges or thinner areas outside theallowable tolerances in the web and flange. Even if this problem isovercome the fabrication, handling, maintenance, and periodicreplacement of the many different cauls would be cost prohibitive.

There thus exists a need to provide a method that controls thethickness, or gage, of the cured laminate material to a substantiallyuniform thickness along the entirety of its curved and noncurved regionsthat overcomes the above problems. These uniformly thick and curvedcomposite parts would find application in a wide variety of differentcommercial applications. For example, these parts would find applicationin the aerospace and automotive industry for parts requiring specificperformance properties and having tight tolerance requirements.

SUMMARY OF THE INVENTION

The present invention provides a method for controlling the thickness ofcomposite laminates cured on closed angel tools. The invention utilizesa peel ply rather than a breather during the preparation for cure oflaminate parts having uniform initial thicknesses. This method allowscomposite parts to be formed with greater uniformity and without theneed for extra hardware.

The composite part formed by the above method may find use in a widevariety of applications, including, for example, automotive andaerospace applications. Thus, for example, one example of a compositepart formed in accordance with the present invention is ideally suitedfor use as a shear tie in a commercial aircraft, which are used tosecure the inner framework of the aircraft to the airplane skin.

Other features, benefits and advantages of the present invention willbecome apparent from the following description of the invention, whenviewed in accordance with the attached drawings and appended claims.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 is a partial view of an aircraft fuselage generally illustratingthe construction thereof.

FIG. 2 is a perspective view of the shear tie of FIG. 1 according to apreferred embodiment of the present invention.

FIG. 3 is another perspective view of the shear tie of FIG. 2.

FIG. 4 is a logic flow diagram for forming the shear ties of FIGS. 1-3.

FIGS. 5-11 are perspective views of various manufacturing steps thatfurther illustrate various portions of the logic flow diagram of FIG. 4as it is used to form the shear tie of FIGS. 2 and 3.

BEST MODES FOR CARRYING OUT THE INVENTION

The present invention provides a method of forming a composite prepregmaterial into a closed angle shape and processing the material in amanner that results in a cured component that is consistent throughoutits thickness. The present invention finds applicable uses in a widevariety of potential applications, including for example, in theaerospace and automotive industry. The preferred method of the presentinvention is ideally suited for forming a composite shear tie 32 that isutilized in the supporting framework of a commercial aircraft. A methodfor forming a shear tie, as one illustrative non-limiting example of apotential end use, will be described further herein below in FIGS. 1-3and 5-11.

Referring to FIGS. 1 and 2, a perspective and partial view of anaircraft fuselage 15, the construction thereof may be observed toconsist of a support frame structure 26 that includes a series of spacedcircumferential frame members 20 that define the overall cross sectionalshape of the aircraft with a series of spaced stringers 22 that runbetween adjacent circumferential frame members 20. Stringers 22 runsubstantially parallel to the longitudinal axis of the aircraft fuselage15 (i.e. from the front of the plane to the back of the plane) while theframe members 20 are substantially transverse to the longitudinal axisof the fuselage 15. An aircraft skin 28 is coupled around support frameassembly structure 26 and is preferably co-cured to the stringers 22. Inaddition, the support frame structure 26 also includes a series ofsupport beams 23 run substantially perpendicular to both the framemembers 20 and stringers 22 and are fastened to the respectivecircumferential frame members 20 using a series of fasteners.

Also shown in FIGS. 1-3 are a series of shear ties 32 that are used tosecure the aircraft skin 28 to the frame members 20 of the fuselage 15.Each shear tie 32 includes a web region 50 and a plurality of curvedflanges 52, otherwise known as tabs 52. Each web region 50 preferablyincludes a series of fastening holes 54. Each tab 52 preferably includesone or more fastening holes 56 and is separated from the next respectivetab 52 by a mouse hole region 58. The web region 50 may optionallyinclude a upwardly extending region 61 along a portion of its length inareas that are to be coupled between the respective frame members 20 andsupport beams 23 and can include a portion of the afore-mentionedfastening holes 54 The fastening holes 54, 56 are introduced to theshear tie 32 in some type of post-production processing well known tothose of ordinary skill in the art.

As best shown in FIG. 3, each tab 52 is bent during processing to arespective angle αwith respect to the web region 50 and therein definesan inner radius 55 between each respective tab 52 and the web region 50.The angle α is set such that the tab 52 substantially abuts itsrespective stringer 22 while the web region 50 substantially abuts itsrespective frame member 20 and is typically about 90 degrees, althoughin alternative embodiments the angle α may be anywhere between about 0degrees and 180 degrees, and more preferably between about 60 and 120degrees. In addition, the thickness t1 of the inner radius 55 within+/−10% of nominal (i.e. the thickness of the part prior to bending),while the thickness t2 of the web region 50 and the thickness t3 of thetab 52 are maintained within +/−6% of nominal.

The shear ties 32 are preferably formed of one or more plies of an epoxyprepreg material that is shaped and curved to conform the outer shape ofthe fuselage 15 and to the respective frame members 20 and stringers 22.Each ply of the epoxy prepreg material consists of carbon fibersimpregnated within an epoxy resin formed by conventional methods wellknown to those of ordinary skill in the art. The epoxy prepreg isusually preformed as a flat layered material and stored in a frozen anduncured state prior to use. The flat prepreg is subsequently thawed,shaped and cured by the preferred method described below generally inFIG. 4, in conjunction with FIGS. 5-11.

As best shown in FIG. 1, each respective shear tie 32 is fastened to arespective frame member 20 through the fastening holes 54 using afastener 62 such that the respective mouse hole region 58 extend overthe respective stringers 22 and such that the curvature of a backsideedge 60 of the web region 50 runs substantially parallel to the innercurve 61 of the frame member 20 between the respective support beams 23.In addition, the shear ties 32 are also preferably fastened between therespective frame members 20 and support beams 23 using one or moreadditional fasteners 67. The tabs 52 are fastened to the airplane skin28 by coupling a fastener 64 through a respective fastening hole 56.Alternatively, the shear tie 32 may also secured to the frame member 20,the support beams 23 and to the skin 28 utilizing an adhesive andwithout the need for fasteners 62, 64 and the respective fastening holes54, 56.

Referring now to FIG. 4, a method for forming a composite material ingeneral, and a shear tie 32 in particular, is described in which thethickness of the radius region is maintained within +/−10% of nominal,while maintaining the non-bent regions within +/−6% of nominal.Maintaining a consistent thickness is important in order to maintain themechanical properties of the composite material, particularly to a sheartie 32 for an aircraft, during subsequent use. Further, maintaining aconsistent thickness is important during assembly fit up to facilitateother parts nesting into this region.

Referring now to Step 200 in FIG. 4, a flat epoxy prepreg member of adesired thickness is formed by conventional methods. Typically, prior touse, the prepreg member 80 is frozen or otherwise maintained in asubstantially uncured state. For a prepreg material (shown as 80 inFIGS. 5-6) that will be used to form a shear tie 32 as shown in FIGS.1-3, the flat epoxy prepreg 80 consists of one of more layers (or“plies”) of carbon fiber impregnated with a 350-degree Fahrenheit curethermosetting epoxy resin.

However, other thermosetting resins may be utilized in forming prepregmembers used in alternative applications. For example, otherthermosetting resins that may be utilized include unsaturated polyesterresins, aminoresins, alkyd resins, phenolic resins, (meth)acrylatedoligomers, silicone resins, and other resins systems that becomethermosetting in the presence of a crosslinking resin such aspolyurethane-isocyanate systems. Further, depending upon the desiredcharacteristics of the laminated material, lower curing temperatureversions of the thermosetting epoxy resin may be utilized such as a250-degree Fahrenheit curing thermosetting epoxy resin. In addition, asone of ordinary skill recognizes, mixtures of the above-listedthermosetting resins may also be utilized.

In addition, other fibrous materials in addition to carbon fibers mayalso be utilized and fall within the spirit of the present invention.For example, glass fibers such as e-type or s-type glass of variouscompositions and radial thicknesses may also be utilized.

Next, in Step 210, the flat epoxy prepreg 80 is cut into desired patternfor manufacture to introduce the various features of the composite partin flat form. For a shear tie 32 as in FIGS. 1-3, this includes the tabs52, the mouse hole regions 58, the web regions 50, as well as shapingcurvatures in the backside 60 and front side 53, as shown best in theirfinal form in FIGS. 2 and 3.

In Step 220, a peel ply layer 82 is applied to one side of the thawedcharge 80 and trimmed to match the contours of the charge 80. To assistin peel ply removal, an insert approved tape tab 84, such as a 1-inchpiece of Teflon tape, is preferably coupled over the peel ply layer 82along the lower edge 60 of the charge 80.

In Step 230, the charge 80 is placed upon two layers of FEP 86 (apolymer film formed of tetrafluoroethylene and hexafluoropropylene) thatare each cut in the approximate shape of the prepreg charge 80 such thatthe prepreg charge 80 is between the peel ply layer 82 and the FEPlayers 86. For a shear tie 32, the mousehole regions 87 are then cut outof the FEP 86, preferably using an X-shape cut 87. Indexing lines 88 aredrawn on the prepreg charge 80 for alignment purposes.

In Step 240, and as shown in FIG. 6, a drape mandrel 90 is placed ontothe prepreg charge 80 such that the peel ply layer 82 contacts themandrel 90 and such that the end 92 of the mandrel 90 is aligned withthe indexing lines 88. As shown in FIG. 7, the prepreg charge 80 is thenheated to a temperature below the curing temperature of the curing resincontained in the prepreg 80 but sufficiently high to allow the warm theresin, allowing the plies of fiber to slip over each other withoutwrinkling and the flat prepreg to be bent to its desired shape over themandrel 90. For a shear tie 32 such as in FIGS. 1-3, the charge isheated to between about 130 and 150 degrees Fahrenheit and bent over thedrape mandrel 90 along the index lines 88, ensuring that the partthickness is maintained within 0.375″ of nominal along its inner radius55.

In Step 250, the bent prepreg charge 80 is removed from the drapemandrel 90 and the FEP layers 86 removed. As shown in FIG. 8, the warmbent prepreg charge 80 is placed in a curing tool 92 such that theradius 55 of the bent prepreg material is in intimate contact with thecure tool 92 and such that the outer portions (here outer portions 57,59, respectively, of the web portion 50 and tabs 52) substantially abutsthe curing tool 92 and such that the peel ply layer 82 does not contactthe cure tool 92 (i.e. is coupled along the side defining the innerradius 55). The prepreg charge 80 is preferably sweeped while the partis warm to ensure good contact of the web portion 50 and tabs 52 to thecuring tool 92.

In Step 260, an optional resin dam 94 is installed along the outerperiphery of the bent charge material 80. For a shear tie 32 as in FIG.8, the resin dam 94 is installed along the inner bend region 53 of thetabs 52 and around the mousehole region 58. Teflon tape is preferablyutilized, taking care to ensure there are no gaps within the tape andminimizing tape bridging, as these bridge areas will fill with resin.Thermocouples (not shown) are preferably installed between the Teflontape and the charge 80.

Next, in Step 270, as also shown in FIG. 8, an optional chromate resindam 98 is applied to the top and bottom of the bent prepreg part 80. Asecond layer of Teflon tape 100 is then installed over the chromateresin dam 98 to avoid sticking to the FEP (shown in FIG. 9 as 104)and/or nylon bag (not shown).

In Step 280, and also as shown in FIG. 9, at least two layers of a 1 milor 2 mil sheet of FEP 104 are installed over the charge in two pieceswith the pieces and overlapping fully one inch along the radius of thepart 80 and 2 inches along the outer edges of the part 80. Further, atleast one layer of the FEP 104 extends beyond the end of the underlyinglayer by at least one-half inch.

In Step 290, and further shown in FIG. 10, a breather 108 is installedaround the part, not to touch the part at any location, offset at leastone inch from the part 80. The breather 108 is continuous from one sideto the other of the part 80 and around the edges of the part 80.

In Step 300, and further shown in FIG. 11, a vacuum bag 110 is installedover the charge 80 and the breather element 108. A pleat 112 is place inthe vacuum bag 110 running the length of the radius of the charge.Preferably, for a shear tie 32, the vertical pleats 112 are introducedin the mousehole region 58 of the charge 80. The charge 80 is then readyfor autoclaving.

Finally, in Step 310, the charge is cured. For the shear tie of FIGS.1-3 and 5-11, the charge 80 is cured using an autoclave. However, forother composite parts, an oven or other heating device may be utilizedto cure the thermosetting resin component of the charge 80.

For an autoclaving process, the breather material 108 and charge 80(laminate) consolidate under heat and pressure. By not placing thebreather element 108 in direct contact with the charge 80, however,material bridging (bagging and part material) in the inner radius 55that typically happens from the breather element being pinned againstthe web portion 50 and flange 52 are eliminated, therein producing apart, here a shear tie 32, with a more consistent thickness throughoutits curved and noncurved regions.

The temperature and pressure used in the autoclave, as one of ordinaryskill recognizes, is dependent upon the curing characteristics of thecurable component used in the charge 80. For a shear tie 32 formed asshown in FIGS. 1-3 and 5-11, utilizing an epoxy prepreg charge 80 havinga 350 degree Fahrenheit curing range preimpregnated with carbon fibers,the process begins by first applying a full vacuum of at least 22 inchesof Mercury. Next, the autoclave is pressurized to at least 85 pounds persquare inch. The autoclave is then ramped up from room temperature to aminimum of 345 degrees Fahrenheit in no more than 5 degrees Fahrenheitper minute increments. The autoclave is then maintained at a minimum of85 pounds per square inch pressure and a temperature between about 345and 365 degrees Fahrenheit for about 120 minutes to ensure completecuring of the 350 degree Fahrenheit curing thermosetting epoxy resin.The autoclave is then cooled to about 140 degrees Fahrenheit in no morethan 5 degrees Fahrenheit per minute increments prior to releasing thepressure and vacuum. The shear tie 32 is then removed from the autoclaveand cooled. The shear tie 32, following further trimming and drilling tofinal part configuration, is then ready to be introduced within thesupport structure 24 of the fuselage 15 as described above.

Tests have confirmed that shear ties 32 formed in accordance with themethod of FIG. 4 achieve a thickness t1 of the inner radius 55 within+/−10% of nominal, while the thickness t2 of the web region 50 and thethickness t3 of the tab 52 are maintained within +/−6% of nominal.

The present invention provides a method of forming a cured compositematerial having a substantially uniform thickness along the entirety ofits curved and noncurved regions with minimal additional steps andwithout the need for extra hardware. The present invention is robust, inthat it can be utilized on composite prepreg materials havingsubstantially varying thicknesses, compositions, and curecharacteristics. Composite parts formed by the present invention areformed with more precision than previously available. Also, thecomposite parts formed according to the present inventions are able tobe used in applications requiring tight clearances. Further, compositeparts having more consistent thickness properties achieve morepredictable mechanical properties. Thus, composite parts formed inaccordance with the present invention may be utilized in a wide varietyof potential applications, including but not limited to aerospace,automotive, and construction applications.

While the invention has been described in terms of preferredembodiments, it will be understood, of course, that the invention is notlimited thereto since modifications may be made by those skilled in theart, particularly in light of the foregoing teachings.

1. A method for forming a composite laminate part having a curved innerradius extending between a pair of outer regions, wherein said curvedinner radius and said pair of outer regions have a substantially uniformthickness after curing as compared with the uncured part, one of thepair of outer regions being bent to an angle α with respect to the otherof the pair of outer regions, the method comprising: providing a flatprepreg charge having a desired shape and a substantially uniforminitial thickness, said flat prepreg material including an uncuredthermosetting polymer preimpregnated with at least one fiber; applying apeel ply layer to one flat side of said flat prepreg charge; couplingone or more layers of FEP to another flat side of said flat prepregcharge such that said flat prepreg charge is contained between said peelply layer and said one or more layers of FEP; coupling said flat prepregcharge to a drape mandrel such that said peel ply layer contacts saiddrape mandrel; bending said flat prepreg charge to the angle α over saiddrape mandrel to form a bent prepreg charge; removing said bent prepregcharge from said drape mandrel; removing said one or more layers of FEPfrom said bent prepreg charge; placing said bent prepreg charge on aremote cure tool such that said peel ply charge does not contact saidfemale cure tool; applying at least two layers of FEP over said bentprepreg charge; coupling a breather element around said outer peripheryof said bent prepreg charge; coupling a vacuum bag over said bentprepreg charge and said breather element; introducing a pleat alongwithin said vacuum bag corresponding to said inner radius of said bentprepreg material; and curing said uncured polymer of said bent prepregmaterial.
 2. The method of claim 1, wherein said uncured thermosettingpolymer comprises a thermosetting epoxy resin.
 3. The method of claim 2,wherein said thermosetting epoxy resin comprises a 350-degree Fahrenheitcurable thermosetting epoxy resin.
 4. The method of claim 1, whereinsaid at least one fiber is selected from the group consisting of carbonfiber and a glass fiber.
 5. The method of claim 4, wherein said glassfiber is selected from the group consisting of e-type glass and s-typeglass.
 6. The method of claim 1, wherein curing said uncured polymer ofsaid bent prepreg material comprises: introducing said bent prepregcharge to an autoclave, applying a full vacuum to said autoclave;pressurizing said autoclave to a desired pressure; heating saidautoclave to a curing temperature; maintaining said autoclave at saidcuring temperature for a period of time sufficient to substantially curesaid uncured polymer; cooling said autoclave to a temperaturesubstantially beneath said curing temperature; depressurizing saidautoclave below said desired pressure; and releasing said full vacuum.7. A shear tie formed in accordance with the method of claim 6, saidshear tie including a web region and a plurality of tabs, wherein eachof said plurality of said tabs is bent to a respective angle α withrespect to said web region and therein defines an inner radius betweeneach respective said tab and said web region, wherein the thickness ofsaid inner radius is within 10% of the thickness of said flat prepregcharge and wherein the thickness of each of said plurality of tabs andsaid web region is maintained within 6% of said substantially uniformthickness of said flat prepreg charge.
 8. An aircraft fuselage includingat least one shear tie formed in accordance with the method of claim 7.9. The method of claim 7, wherein said respective angle α is between 60and 120 degrees.
 10. The method of claim 7, wherein said respectiveangle α is about 90 degrees.
 11. The method of claim 1, wherein bendingsaid flat prepreg charge over said drape mandrel to form a bent prepregcharge comprises: heating said flat prepreg charge to a temperaturesufficient to bend said flat prepreg charge to a desired angle, whereinsaid temperature is less than the curing temperature of said uncuredthermosetting polymer; and bending a portion of said heated flat prepregcharge over said drape mandrel.
 12. The method of claim 7, whereincuring said uncured polymer of said bent prepreg material comprises:introducing said bent prepreg charge to an autoclave, said bent prepregcomprising a 350 degree Fahrenheit curing thermosetting epoxy resinimpregnated with at least one carbon fiber, applying a full vacuum of atleast 12 inches of mercury to said autoclave; pressurizing saidautoclave to at least 85 pounds per square inch; heating said autoclaveto at least 345 degrees Fahrenheit in increments not to exceed 5 degreesFahrenheit per minute; maintaining said autoclave at between 345 and 365degrees Fahrenheit for about 120 minutes to substantially cure saiduncured polymer; cooling said autoclave to about 140 degrees Fahrenheitin increments not to exceed 5 degrees Fahrenheit per minute;depressurizing said autoclave from at least 85 pounds per square inch toan ambient pressure; and releasing said full vacuum.
 13. A method forforming a composite laminate having a substantially uniform thicknessthroughout its curved and uncurved regions from a preformed bent anduncured prepreg charge, the composite laminate having the same relativethickness post curing as it had pre-curing, the method comprising:providing the preformed bent prepreg charge having a curved inner radiusand a pair of outer regions extending from said curved inner radius,said one of said pair of outer regions being bent to an angle α withrespect to said other of said pair of outer regions, said curved innerradius and said pair of outer regions having substantially the samethickness, said bent prepreg material including an uncured thermosettingpolymer preimpregnated with at least one fiber; coupling a peel plylayer along a first side of said bent prepreg charge such that said peelply layer substantially contacts an inner portion of said curved innerradius; placing said bent prepreg charge on a remote cure tool such thatsaid peel ply charge does not contact said female cure tool; applying alayer of FEP over said bent prepreg charge; coupling a breather elementaround said outer periphery of said bent prepreg charge; coupling avacuum bag over said bent prepreg charge and said breather element;introducing a pleat along within said vacuum bag corresponding to saidinner radius of said bent prepreg material; and curing said uncuredpolymer of said bent prepreg material.
 14. The method of claim 13,wherein said uncured thermosetting polymer comprises a thermosettingepoxy resin.
 15. The method of claim 14, wherein said thermosettingepoxy resin comprises a 350 degree Fahrenheit curable thermosettingepoxy resin.
 16. The method of claim 13, wherein said at least one fiberis selected from the group consisting of carbon fiber material and glassfiber material.
 17. The method of claim 16, wherein said glass fiber isselected from the group consisting of e-type glass and s-type glass. 18.A shear tie formed in accordance with the method of claim 12, said sheartie including a web region and a plurality of tabs, wherein each of saidplurality of said tabs is bent to a respective angle α with respect tosaid web region and therein defines a curved inner radius between eachrespective said tab and said web region, wherein the thickness of saidcurved inner radius is within 10% of the thickness of said bent prepregcharge and wherein the thickness of each of said plurality of tabs andsaid web region is maintained within 6% of said substantially uniformthickness of said bent prepreg charge.
 19. An aircraft fuselageincluding at least one shear tie formed in accordance with the method ofclaim
 18. 20. The method of claim 18, wherein curing said uncuredpolymer of said bent prepreg material comprises: introducing said bentprepreg charge to an autoclave, said bent prepreg charge comprising a350 degree Fahrenheit curing thermosetting epoxy resin impregnated withat least one carbon fiber, applying a full vacuum of at least 22 inchesof mercury to said autoclave; pressurizing said autoclave to at least 85pounds per square inch; heating said autoclave to at least 345 degreesFahrenheit in increments not to exceed 5 degrees Fahrenheit per minute;maintaining said autoclave at between 345 and 365 degrees Fahrenheit forabout 120 minutes to substantially cure said uncured polymer; coolingsaid autoclave to about 140 degrees Fahrenheit in increments not toexceed 5 degrees Fahrenheit per minute; depressurizing said autoclavefrom at least 85 pounds per square inch to an ambient pressure; andreleasing said full vacuum.